Vertical take off aircraft

ABSTRACT

An aircraft includes a fuselage, a wing, a ducted fan and a controller. The wing and the ducted fan are coupled to the fuselage. The controller is operable to control the aircraft in a vertical flight mode, a horizontal flight mode, and transition the aircraft from the vertical flight mode to the horizontal flight mode.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.15/402,784, filed on Jan. 10, 2017, which is a continuation of U.S.patent application Ser. No. 15/152,413 filed on May 11, 2016, now issuedas U.S. Pat. No. 9,828,090 on Nov. 28, 2017, which is a continuation ofU.S. patent application Ser. No. 14/837,866, filed on Aug. 27, 2015 nowissued as U.S. Pat. No. 9,365,290. This application is a continuation ofU.S. patent application Ser. No. 15/402,784, filed on Jan. 10, 2017,which is also a continuation of U.S. patent application Ser. No.15/152,416 filed on May 11, 2016, now issued as U.S. Pat. No. 9,567,071on Feb. 14, 2017, which is a continuation of U.S. patent applicationSer. No. 14/837,866, filed on Aug. 27, 2015 now issued as U.S. Pat. No.9,365,290. The contents of which are hereby incorporated by reference intheir entirety for all purposes.

BACKGROUND

Aircraft are generally considered to be one of two different types basedon how they generate lift. A first type generates lift force from airpassing over a fixed wing. This type of aircraft generally has a goodlifting capacity, which also means that it can carry large amounts offuel for sustained flight. However, because air needs to pass over thewing to generate the lift force, the aircraft must be accelerated on theground until sufficient lift can be generated for take off. This limitsthe use of the aircraft to locations where there are suitable andaccessible areas to accommodate take off and landing procedure.

A second type of air craft generates lift force from a rotating blade.Because the blade itself rotates, a lift force can be generated whilethe aircraft is stationary on the ground enabling it to be deployed fromlocations that cannot accommodate take off and landing of a fixed wingaircraft. However, significantly more power is required to generate alifting force thereby limiting the fuel capacity and range of theaircraft. This type of aircraft is also generally limited to lower airspeeds.

There have been some attempts to provide aircraft having a fixed wingand vertical take off capabilities. For example, the F-35 has a verticaltake off variant. But, to provide this capability uses an extremelycomplex reconfiguration of the aircraft. In the F-35, the thrust isvectored toward the ground and a bay having a vertical fan is openednear the front of the aircraft. This reconfiguring is mechanicallycomplex and also makes transitioning the aircraft extremely complex.This type of configuration generally also requires a significant amountof power to lift the aircraft such that the payload and even fuelcapacity of the aircraft is reduced limiting its range.

It would be desirable to have an aircraft that combined the range andspeed capabilities of the fixed wing aircraft with the take off andlanding capabilities of a rotary blade aircraft. It would also bedesirable to have such an aircraft of small or medium size that could beeasily transported, deployed in a wide variety of terrain, and remotelyoperated or autonomously operated.

BRIEF SUMMARY

In an embodiment, an aircraft includes a fuselage, a wing, a ducted fan,a propulsion unit, and a controller. The fuselage has a nose end and asecond, opposing end defined as a tail end. The wing coupled to thefuselage between the nose end and the tail end. The ducted fan iscoupled to the fuselage at a point between a location where the wing iscoupled to the fuselage and the tail end of the fuselage. The propulsionunit is coupled to the ducted fan and is operable with the ducted fan togenerate a thrust at least equal to a weight of the aircraft. Thecontroller is operable to control the aircraft in (1) a vertical flightmode in which a first lifting force generated by the ducted fan isgreater than a second lifting force generated by the wing, (2) ahorizontal flight mode in which the second lifting force is greater thanthe first lifting force, and (3) transition the aircraft from thevertical flight mode to the horizontal flight mode.

In an embodiment, an aircraft includes a ducted fan, a fuselage, a wing,and a controller. The ducted fan has a plurality of controllable vanes.The fuselage is disposed at a first side of the ducted fan and coupledto the ducted fan via one or more connecting members The wing is coupledto the fuselage and disposed at the first side of the ducted fan. Thecontroller is operable to control the aircraft in a vertical flightmode, a horizontal flight mode, and transition the aircraft between thevertical flight mode and the horizontal flight mode.

In a embodiment, a method of transitioning an aircraft that includes aducted fan having a plurality of radial controllable vanes, a fuselagedisposed at a first side of the ducted fan and coupled to the ducted fanvia one or more struts, and a wing coupled to the fuselage and disposedat the first side of the ducted fan, includes: flying the aircraft in avertical flight mode in which a first lifting force generated by theducted fan is greater than a second lifting force generated by the wing;adjusting the output power and pitching moment of the controllable vanesto accelerate the aircraft in a horizontal direction; and flying theaircraft in a horizontal flight mode in which the second lifting forceis greater than the first lifting force.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of an aircraft according to an embodiment.

FIG. 2 is a side view of the aircraft of FIG. 1.

FIG. 3 is another side view of the aircraft of FIG. 1.

FIG. 4 is a top view of the aircraft of FIG. 1.

FIG. 5 is a plan view of exemplary control hardware in accordance withthe disclosure.

FIG. 6A is a perspective illustration of an aircraft coordinate system.

FIG. 6B is a perspective illustration of an earth coordinate system.

FIG. 7 is a tail forward view of an exemplary ducted fan.

FIG. 8A is a tail forward view of an exemplary ducted fan in a positiveroll configuration.

FIG. 8B is a tail forward view of an exemplary ducted fan in a positivepitch configuration.

FIG. 8C is a tail forward view of an exemplary ducted fan in a positiveyaw configuration.

FIG. 9 is a flow diagram of an exemplary outbound transition.

FIG. 10 is a block diagram of an exemplary control algorithm forgenerating Euler angle commands.

FIG. 11 is a block diagram of an exemplary control algorithm forgenerating Euler angle errors.

FIG. 12 is a block diagram of an exemplary control algorithm forgenerating body axis Euler angles.

FIG. 13 is a block diagram of an exemplary control algorithm forgenerating roll, pitch and yaw commands.

FIG. 14A is a block diagram of an exemplary control algorithm fordetermining vane deflections.

FIG. 14B is a tail forward view of an exemplary ducted fan illustratingindexing that may be used with the block diagram of FIG. 14A.

FIG. 14C is a block diagram of an exemplary control algorithm for analtitude controller.

DETAILED DESCRIPTION

Various aspects of an aircraft according to the present disclosure aredescribed. It is to be understood, however, that the followingexplanation is merely exemplary in describing the devices and methods ofthe present disclosure. Accordingly, any number of reasonable andforeseeable modifications, changes, and/or substitutions arecontemplated without departing from the spirit and scope of the presentdisclosure.

Referring to FIGS. 1-4, an aircraft 100 includes a fuselage 102, a wing104, and a ducted fan lift system 106 with integral control vanes 108.The ducted fan lift system 106 may include: a duct 110 that is coupledto the fuselage 102 by struts 112, and ground contact protrusions 114extending from the duct 110. The ducted fan system 106 provides enoughthrust at zero velocity to allow the aircraft to hover, preferably 25%more thrust than the maximum vehicle weight.

The duct 110 is located at a distance below the wing 104 (i.e., the wing104 is disposed between a nose of the aircraft and the ducted fan system106) such that the aircraft's center of gravity (CG) 114 provideslateral and longitudinal stability in wing-borne flight as well ascontrol authority in hover. It will be appreciated that the distancebetween a tail 118 of the aircraft and a center axis C of the wing maybe determined by theoretical analysis and flight tests. In an example,the CG of the aircraft may be disposed at a longitudinal position underthe wing of the aircraft. The duct 110 and the control vanes 108 producerestoring moments to align the aircraft to the relative wind, similar tothe tail of a conventional aircraft. The control vanes 108 may bedeflected to change the moment on the aircraft 100 (as measured aboutthe CG) and they have limited control power to do so based on theirsize, location relative to the CG, location in the duct, and the airflowthrough the duct. The duct 110 and vanes 108 are preferably far enoughbelow the CG to provide stability (e.g., restoring moment, which isimportant for stable high speed flight) and for control power (e.g.,ability to create a control moment by deflecting the jet of airemanating from the duct). By analyzing the complete aerodynamics of theaircraft and testing it for all points (speeds) in the transitionmaneuver, the inventors discovered that it is possible to balance thesemoments using vane deflection that does not cause the vanes to block theairflow in the duct, for example, within 30 degrees from neutral.

Using a ducted fan for propulsion may provide for the exemplaryadvantages of: 1) protecting people from the fan blades; 2) serving as astabilizing surface (e.g., it may replace tail surfaces) when mounted atthe back of the aircraft; 3) augmenting the thrust of the fan for agiven amount of power input; and 4) allowing for the use of controlvanes disposed in the high velocity flow of the fan, which may improvethe control authority (for example, as compared to a non-ducted designor a design that might have the propeller mounted on the front of thefuselage).

The control vanes 108 inside the duct are arranged and sized (vanenumber, geometry, and area) to provide control authority (roll, pitchand yaw torque) in both hover and wing-borne flight. The desired controlauthority implies a minimum disc loading (i.e. the vehicle's weightdivided by the internal disc area) to provide enough velocity over thecontrol vanes 108.

The wing area of the wing 104 is preferably chosen to support theaircraft's weight at a designated speed for the mission. The wing areamay be increased from the size selected for cruise on a similar aircraftthat operates only in a horizontal flight mode, to facilitate transitionfrom vertical to horizontal flight. For example, the wing area in ahorizontal only flight mode aircraft may be selected so that at cruisespeed it provides lift equal to the aircrafts maximum weight. If therewere no wing, the aircraft would be limited in its maximum tilt angle toa value where the peak thrust time the cosine of the tilt angle is equalto the weight of the aircraft. If the tilt angle exceeds this amount,the aircraft will lose altitude because the vertical thrust will be lessthan the weight of the aircraft. Inclusion of the wing allows theaircraft to balance its weight at greater tilt angles because the wingwill begin to carry the weight (e.g., air begins flowing over the wingand the wing generates lifting force) as the aircraft tilts and gainsspeed. The wing size (as well as the other design variables) may beselected so that the aircraft may balance the moments (roll, pitch, yaw)and forces (lift, weight, thrust, drag) at each airspeed during thetransition maneuver.

The center of gravity 114 of the aircraft may be located at a fractionof the wing chord selected for static pitch stability. The use of afeedback control system allows for more flexibility in the location ofthe center of gravity. Moving the wing further forward from the CG maydestabilize the aircraft in high speed flight but may also produce amoment that helps balance the restoring force of the duct below the CG,making it possible to trim the aircraft with less vane deflection.Preferably, the wing size and distance from the duct, and CG locationare selected to provide static stability in high speed flight.

The dimensions and location of the wing of the aircraft 100 arepreferably selected to balance the forces and moments acting on theaircraft at any speed between hover and wing-borne flight. This mayfacilitate the transition maneuver described below.

Thus, a tail-sitter ducted fan VTOL (vertical take off and landing)aircraft may be capable of transitioning with reduced altitudedeviations (e.g., less than 100 ft through the transition). The designvariables that may be selected in accordance with this disclosureinclude engine power (e.g., installed maximum horsepower), duct innerdiameter, duct chord (or height), duct shape (e.g., airfoil crosssection), vane size (e.g., chord), vane shape (which may includeadditional surfaces at vane tips), vane number, vane arrangement/layout(e.g., radial arrangement, cross arrangement, etc.), vane positionrelative to duct and aircraft CG, stator size (e.g., the surfaces thatlie in front of the vanes), stator shape, aircraft CG location, wingspan, wing area, and wing location relative to aircraft CG.

Exemplary design constraints include: 1) roll, pitch, and yaw torquesare preferably balanced throughout the transition flight envelope (e.g.,control authority from the vanes and any other controls); 2) vanedeflection is preferably below the vane saturation limit duringtransition (the saturation limit is a combined property of the designvariables and occurs when the vanes are deflected to the point where amaximum control effect is reached; 3) forces on the vehicle arepreferably balanced at all phases of the transition maneuver so thataltitude does not change significantly and airspeed can be regulatedthroughout the maneuver (engine power and wing area are significantdesign variables with respect to this constraint); 4) the aircraft ispreferably statically stable in pitch and yaw when in the high speedflight mode (CG location and duct size relative to the wing size aresignificant variables with respect to this constraint); and 5) theaircraft preferably meets mission performance requirements (such asrange, duration, ceiling, speed, payload capability; wing area, ductsize, engine performance, and structural weight are significant designvariables with respect to this constraint).

In a particular example, the fuselage may have a width FW of 8 inches, aheight FH of 7 inches and a length FL of 66.5 inches. The nose cone ofthe aircraft may have a length of 8.75 inches. The center chord C of thewing 104 may be 70 inches from the nose of the aircraft. The aircraftmay have a width AW of 108 inches, which may also be the wing span, anda length AL of 98.25 inches. The wing area may be 1058 in². The enginepower may be 13 HP. The duct may have eight vanes in a radialarrangement. The stators may be arranged radially and located ahead ofeach vane. Each stator may have a chord of 1.75″ and a thickness of0.7″. The leading edge may have a circular radius of 0.35″ and thetrailing edge of the stator may be blunt and resting against the leadingedge of each vane. The stators may be angled to align with the swirlvelocity of the flow from the fan (or propeller).

Control Hardware

FIG. 5 is a plan view of exemplary hardware in accordance with thedisclosure. Aviation control module 200 is an example of specializedhardware for controlling the aircraft. The aviation control module 200may include an embedded processor, memory (both volatile andnon-volatile) that may be programmed to carryout the control algorithmsdiscussed in this disclosure. The aviation control module 200 includesanalog and digital input circuits for receiving information from sensorsonboard the aircraft. The received information, such as airspeed,orientation and location of the aircraft, may be used by the aviationcontrol module to carry out the control algorithms. It will beappreciated that sensors (and actuators) coupled to the aviation controlmodule may alternatively be provided coupled to other module (or controlmodules) in communication with the aviation control module 200 by a databus. Examples of sensors in communication with the aviation controlmodule 200 include air speed sensor 202 (such as a pitot static airtube), and accelerometer or gyro sensor 204 that may sense theorientation of the aircraft in one or more axes. A GPS receiver may beintegrated with the aviation control module 200 or also provided as anadditional module. The aviation control module 200 may also be coupledto an antenna 206 for use by the GPS receiver and/or communications withthe ground or other aircraft. The aviation control module 200 may alsobe connected to actuators 208 a and 208 b associated with left and rightailerons as well as the control vane actuators via the control bus 210and the engine throttle control (e.g., for an internal combustionengine, a turbine, or an electric motor coupled to the vanes of theducted fan) via the signal wires (or control bus) 212. The combinationsof the sensors, the actuators and the specially configured aviationcontrol module provide a specialized and improved platform directed tothe technology of the aircraft.

Coordinate System

The following discussion refers to two coordinate systems, which areright handed orthogonal system. FIG. 6A illustrates an aircraftcoordinate system. The aircraft coordinate system includes the bodyaxes: X body=nose forward; Y body=right wing; and Z body=towardsfuselage bottom (belly). FIG. 6B illustrates an earth coordinate system.The earth coordinate system includes the earth axes: X earth=North; Yearth=East; and Z earth=Down.

Euler Angles

Euler angles are an ordered set of rotations about the axes definedabove that describe a change in orientation from one frame of referenceto another. The rotation order is important in these definitions. Hoverand transition Euler angles include a transform from inertial (earth) tobody axes using the rotation order: psi (z down, earth), phi (x noseforward, prime), theta (y right wing, body). High Speed Euler Anglesinclude a transform from inertial to body axes, using the rotationorder: psi (z down, earth), theta (y right wing, prime), phi (x noseforward, body).

It will be appreciated that the discussion of Euler angles is exemplaryand other techniques for describing rotations between coordinate systems(such as quaternions) are within the scope of this disclosure.

Control Axes

In the following discussion, positive (+) roll is defined as right wingdown along the body x axis; positive (+) pitch is defined as nose upalong the body y axis; and positive (+) yaw is defined as nose rightalong the body z axis.

Control Deflection

The aircraft may include 8 control vanes 108 located in the duct 110.The vanes 108 may be distributed evenly every 45 degrees along the radiiof the duct 110, as shown in FIG. 7. The vanes may be controlled bythree high level commands: roll, pitch, and yaw. Each command may berealized as an instruction to deflect the air moving through the duct ina direction that creates a moment about the aircraft's center of gravitythat produces the commanded effect. The mapping of the vane deflectionsis provided below where i is an index of the vane and δ is the magnitudeof the deflection, and Φ is the roll associated with that vane. Apositive vane deflection is right hand rule positive outward along theradii of the vane.Roll(i)=−δ rollPitch(i)=−δ pitch*sin(Φ_(vane)(i))Yaw(i)=δ yaw*cos (Φ_(vane)(i))

The total deflection for each vane is the sum of each command definedabove:Vane(i)=Roll(i)+Pitch(i)+Yaw(i)

FIG. 8A illustrates a vane configuration for a positive roll. FIG. 8Billustrates a vane configuration for a positive pitch. FIG. 8Cillustrates a vane configuration for a positive yaw.

When ailerons contribute to the control moment, that contribution isrepresented by δ roll.

Outbound Transition

Outbound transition may refer to the portion of the flight where theaircraft starts in a hover (e.g., vertical flight), for example at acertain height and location, and then accelerates, preferably in asmooth and stable manner, to a speed where the aircraft's wing willsupport the entire weight of the vehicle. The position, orientation(e.g., Euler angles for hover), body axis rotation rates, airspeed,altitude, ground speed, and velocity heading may all be sensed orestimated by the control algorithm for outbound transition of the onboard flight control system at a frequency sufficient for controllingthe vehicle. A wind estimator may be used to orient the bottom of thefuselage towards the wind prior to transition, and the wind directionmay define the courseline heading along which the transition will takeplace. Orienting the aircraft into the wind may provide for a quickertransition as the aircraft will tilt into the wind to maintain a hoverat a specified location. Thus, an air flow direction perpendicular tothe span of the wing improves the lift force that is generated. However,it will be appreciated that the transition may also be performed alongany heading.

During the transition, the wing may be directed into the wind (e.g., 90degrees into the wind). This orientation may create more drag than, forexample aligning the wing to the air flow but it will also provide liftgenerated by the wing more quickly and therefore reduce the altitudeloss in transition. It will be appreciated that this wing alignmentapproach is not limited to an aircraft with a ducted fan at the tail andmay also be used with aircrafts having propellers and ducted fans atother locations on the aircraft.

The control algorithm may be initialized with the location, heading, andaltitude for the beginning of the transition maneuver. Then, theaircraft's ground speed may be commanded to increase at a prescribedacceleration rate (e.g., by a trajectory generator) until the sensedairspeed exceeds the stall speed by a specified margin. At this pointthe transition may be considered complete and control may be turned overto a high speed flight control algorithm. The flight mode commands fromthe ground station operator may be associated with a mission flight planincluding a list of waypoints, speeds, altitudes, and desired flightmodes. Direct commands to the aircraft may also be used. During thetransition maneuver the control system preferably maintains altitude andcourseline heading while following the commanded acceleration profilealong the commanded trajectory.

FIG. 9 illustrates an exemplary outbound transition. At step S1, sensoryinputs such as speed, position and altitude and trajectory commands suchas acceleration and courseline heading are used to generate Euler anglecommands for pitch, roll and yaw (e.g., ψ_(com), ϕ_(com), θ_(com)respectively). For example, Euler commands may be generated to orientthe aircraft in a prescribed heading and at a prescribed speed to carryout the transition by comparing the sensor inputs to the desiredorientation of the aircraft based on the trajectory commands. A specificexample of the control logic is provided in FIG. 10.

At step S3, Euler angle errors (e.g., ψ_(err), ϕ_(err), θ_(err)) may begenerated by comparing the Euler angle commands from step S1 with Eulerangles associated with the current orientation of the aircraft, forexample as shown in FIG. 11.

At step S5, the Euler angle errors may be converted to body axis errors(e.g., ψ_(body_err), ϕ_(body_err), θ_(body_err)) using the conversionsdiscussed above. An example of the control logic to convert the errorsis provided in FIG. 12.

At step S7, roll, pitch and yaw commands are determined to generate vanecommands (e.g., δ_(roll), δ_(pitch), δ_(yaw)). An example of the controllogic to convert the errors is provided in FIG. 13. The variables usedin FIG. 13 are defined as follows:

p=roll rate (about x-body axis)

pdot=roll acceleration (about x-body axis)

ϕbody error=computed roll angle (about x-body axis)

δroll=roll axis vane command (used to compute individual vanedeflections)

q=pitch rate (about y-body axis)

qdot=pitch acceleration (about y-body axis)

θbody error=computed pitch angle (about y-body axis)

δpitch =pitch axis vane command (used to compute individual vanedeflections)

r=yaw rate (about z-body axis)

rdot=yaw acceleration (about z-body axis)

ψbody error=computed yaw angle (about z-body axis)

δyaw=yaw axis vane command (used to compute individual vane deflections)

K8-16=control gains

At step S9, the vane commands are converted to vane deflections. Forexample, if the vane command is for 20% roll, 30% pitch and 40% yaw andthe deflections in FIGS. 8A-8C represent 100% roll, pitch and yaw, theneach vane may be assigned a deflection associated with 20% of itsposition in FIG. 8A, 30% of its position in FIG. 8B and 40% of itsposition in FIG. 8C. Another example of a control algorithm to determinevane deflection is shown in FIG. 14A. FIG. 14B illustrates an indexingpattern for the vanes that may be used with the control algorithm ofFIG. 14A. FIG. 14C illustrates an exemplary control algorithm for analtitude controller to determine a throttle setting. The variables usedin FIG. 14 are defined as follows:

δroll=roll axis vane command (used to compute individual vanedeflections)

δpitch=pitch axis vane command (used to compute individual vanedeflections)

δyaw=yaw axis vane command (used to compute individual vane deflections)

γvane(i)=individual vane azimuth angle as shown in FIG. 14B.

δvane(i)=ith computed vane deflection

δailerons=computed aileron deflection

δthrottle=yaw computed engine throttle setting

h=altitude

hcommand=altitude command

herr=altitude error

hdot=altitude rate

At step S11, the air speed of the aircraft is checked to see if it isabove a threshold (e.g., 125% of the stall speed). If so, the processadvances to the next flight mode (e.g., high speed flight) at step S13.If not, the process returned to step S3 and thereafter repeats until thetransition is completed.

The outbound transition flight control block diagram may include lowlevel control loops that determine how the control surfaces and throttlewill be commanded to achieve the desired motion. There may be fourcontrol feedback loops each commanding a control axis (δroll, δpitch,δyaw) based on a PID (proportional integral derivative) control logic.The variables being regulated may be the body axis orientation commandsrequired to achieve the trajectory guidance previously described as wellas speed and altitude. Of course, it will be appreciated that othercontrol algorithms and variations to the disclosed algorithms are alsocontemplated.

The body axis rotation angles may be calculated at a level above the lowlevel PID feedback control loop. These angles represent the totalrotation about each of the vehicle's orthogonal body axes (see FIG. 6)that may align the aircraft's velocity vector with the computedcourseline for the transition maneuver as well as maintain the desiredacceleration profile. FIG. 12 shows the determination of the body axisrotation angles. First, rotation commands may be determined that maycause the vehicle's orientation (as defined by the Euler angles used inhover) to align with the prescribed transition trajectory and speedcommand. These commands may be in a non-orthogonal set of axes that arenot fixed in the body of the aircraft. Kinematic relations may be usedto transform each command Euler rotation into the orthogonal vehiclebody axis coordinate system so that they can be used for direct feedbackcontrol of the roll, pitch, and yaw commands. These body axis rotationcommands may also be based on PID control logic so that the desiredvariables can be regulated in a robust manner.

The velocity error between the vehicle's ground speed and the commandedspeed may affect the pitch angle command and the cross track error fromthe desired courseline may affect the bank angle command. The headingcommand may be determined by the error between the aircraft's actualheading and the heading of the specified transition courseline, whichmay be a function of the local wind direction and terrainconsiderations. The altitude error between the commanded and actualaltitude may be used to command the throttle control.

FIGS. 10-13 shows how the Euler angle commands may be generated and thentransformed into the body axis rotation angles which may then used inthe inner control loops to determine the roll, pitch, and yaw commands.FIG. 14 shows how the roll, pitch, and yaw commands may be used tocalculate the deflection of each control vane.

With respect to FIGS. 10-14, the depicted control algorithms include thepossibility for gain scaling (coefficients K_(x)) based on airspeed andother parameters. The gain scaling for airspeed may increase stabilityas the airspeed increases from hover through transition and themagnitude of each gain may be adjusted based on the instantaneousairspeed of the vehicle.

These control loops may be executed as described until the prescribedtransition airspeed is exceeded (e.g., 125% of the stall speed) at whichpoint the flight mode is switched to the high speed mode or to the nextcommanded value if the aircraft receives a direct command from theoperator.

Inbound Transition

Inbound transition refers to the portion of the flight where theaircraft starts in an established high speed flight (e.g., wing-borneflight) at a height, airspeed, and location and then decelerates,preferably in a smooth and stable manner, to a speed where theaircraft's motion over the ground is lower than a specified ground speedvelocity threshold and then switches to a hover mode controller. Thecontrol algorithm for inbound transition may be the same as the onedescribed above for the outbound transition except that the initialconditions differ and the velocity is commanded at a specifieddeceleration rate rather than accelerating. A wind estimator may be usedto orient the bottom of the fuselage towards the wind prior to inboundtransition. The wind direction may preferably define the courselineheading along which the transition will take place. The position wherethe transition will be conducted can be estimated prior to the maneuverand the starting point can then be chosen so that the maneuver will becompleted near a specified location, such as the landing zone. Thecontrol algorithm may be initialized with the location, heading, groundspeed, and altitude for the beginning of the inbound transitionmaneuver. The control algorithm may also be initialized with thelocation, heading, ground speed, and altitude for the end of the inboundtransition maneuver and the values for the beginning of the maneuvercalculated from the desired ending values. The maneuver may beconsidered complete when the actual groundspeed is lower than theprescribed hover threshold speed and then the flight mode may beswitched to hover unless commanded otherwise by the flight plan orground station operator.

The commanded deceleration rate is preferably chosen to avoidsignificant altitude increase during the inbound transition (e.g., lessthan 100 ft altitude variation through the transition). If thedeceleration command is too rapid the vehicle may climb excessively evenat the minimum allowable throttle setting and therefore take longer tohover back down to the landing altitude. The aircraft may still beoperated safely with such a flare, but is it preferable to select a lowdeceleration rate to reduce or prevent this from happening.

Another approach to reducing excessive climbing during the inboundtransition is to include control surfaces on the wing that diminish thelift (e.g. “spoilers”) during the inbound transition and thereby makethe maneuver less sensitive to the commanded deceleration rate. Thesecontrol surfaces may be deployed by the feedback control system toreduce lift as needed to maintain the commanded height.

The breadth and scope of the present disclosure should not be limited byany of the above-described exemplary embodiments, but should be definedonly in accordance with the following claims and their equivalents.Moreover, the above advantages and features are provided in describedembodiments, but shall not limit the application of the claims toprocesses and structures accomplishing any or all of the aboveadvantages.

Additionally, the section headings herein are provided for consistencywith the suggestions under 37 CFR 1.77 or otherwise to provideorganizational cues. These headings shall not limit or characterize theinvention(s) set out in any claims that may issue from this disclosure.Specifically and by way of example, the claims should not be limited bythe language chosen under a heading to describe the so-called technicalfield. Further, a description of a technology in the “Background” is notto be construed as an admission that technology is prior art to anyinvention(s) in this disclosure. Neither is the “Brief Summary” to beconsidered as a characterization of the invention(s) set forth in theclaims found herein. Furthermore, any reference in this disclosure to“invention” in the singular should not be used to argue that there isonly a single point of novelty claimed in this disclosure. Multipleinventions may be set forth according to the limitations of the multipleclaims associated with this disclosure, and the claims accordinglydefine the invention(s), and their equivalents, that are protectedthereby. In all instances, the scope of the claims shall be consideredon their own merits in light of the specification, but should not beconstrained by the headings set forth herein.

What is claimed is:
 1. An aircraft, comprising: a fuselage having a noseend and a second, opposing end defined as a tail end; a wing coupled tothe fuselage between the nose end and the tail end; a propulsion unitcoupled to the fuselage and disposed proximal the tail end; and acontroller configured to control the aircraft in: a vertical flight modein which a first lifting force generated by the propulsion unit isgreater than a second lifting force generated by the wing, a horizontalflight mode in which the second lifting force is greater than the firstlifting force, and a transition flight mode in which the aircrafttransitions from the vertical flight mode to the horizontal flight mode,wherein in the transition flight mode, the controller is operable to:accelerate the aircraft until a sensed airspeed exceeds a threshold, andorient the aircraft based on trajectory commands by comparing sensorinputs to a desired orientation.
 2. The aircraft of claim 1, wherein thepropulsion unit is configured to provide directed thrust.
 3. Theaircraft of claim 1, wherein, in the transition flight mode, theaircraft is configured to determine a ground speed acceleration rate andaccelerate the aircraft at the ground speed acceleration rate.
 4. Theaircraft of claim 1, wherein the threshold is above the stall speed by amargin.
 5. The aircraft of claim 1, wherein the propulsion unit includesa ducted fan.
 6. The aircraft of claim 5, wherein the propulsion unitincludes a plurality of control vanes to provide directed thrust.
 7. Theaircraft of claim 1, wherein the controller is configured to generatecommands for pitch, speed and roll based at least in part on sensoryinput values for speed, position and altitude, and determine Euler angleerrors based on the sensory input values and the trajectory commands. 8.The aircraft of claim 7, wherein the controller is configured todetermine body axis angle errors based on the Euler angle errors.
 9. Theaircraft of claim 1, wherein the controller is configured to transitionthe aircraft from the vertical flight mode to the horizontal flight modewhile maintaining altitude of the aircraft within 100 ft throughout thetransition.
 10. The aircraft of claim 1, wherein the controller isconfigured to transition the aircraft from the horizontal flight mode tothe vertical flight mode.
 11. The aircraft of claim 10, wherein thecontroller is configured to transition the aircraft from the horizontalflight mode to the vertical flight mode while maintaining altitude ofthe aircraft within 100 ft throughout the transition.
 12. The aircraftof claim 1, wherein a center of gravity of the aircraft is disposed in alocation such that the aircraft balances at rest in an orientation wherethe nose end is positioned over the duct.
 13. A method of transitioningan aircraft that includes a fuselage, a propulsion unit and a wing,wherein the fuselage includes a nose end and a second, opposing enddefined as a tail end, the propulsion unit is coupled to the fuselageand disposed proximal the tail end, and the wing is coupled to thefuselage, the method comprising: flying the aircraft in a verticalflight mode in which a first lifting force generated by the propulsionunit is greater than a second lifting force generated by the wing;adjusting an output power of the propulsion unit and a pitching momentof the aircraft to accelerate the aircraft in a horizontal direction,the adjusting including accelerating the aircraft until a sensedairspeed exceeds a threshold; flying the aircraft in a horizontal flightmode in which the second lifting force is greater than the first liftingforce; and flying the aircraft at an orientation based on trajectorycommands by comparing sensor inputs to a desired orientation.
 14. Themethod of claim 13, wherein the propulsion unit includes a ducted fan.15. The method of claim 13, wherein the propulsion unit includes aplurality of control vanes to provide directed thrust, and the adjustingincludes adjusting a position of the control vanes.
 16. The method ofclaim 13, further comprising determining Euler angle errors based onsensory input values and the trajectory commands.
 17. The method ofclaim 16, further comprising determining body axis angle errors based onthe Euler angle errors.
 18. The method of claim 17, wherein the flyingthe aircraft in the vertical flight mode, the adjusting, and the flyingthe aircraft in the horizontal mode are performed maintaining altitudeof the aircraft within 100 feet.
 19. The method of claim 13, furthercomprising adjusting, after the flying the aircraft in a horizontalflight, the output power and pitch moment of the controllable vanes todecelerate the aircraft in the horizontal direction until the firstlifting force is greater than the second lifting force.
 20. The methodof claim 13, further comprising maintaining trim of the aircraft at anyangle of attack between zero degrees and ninety degrees.